The resultant force on a wing section can be specified by two components of force perpendicular and parallel to the air stream (the lift and drag respectively) and by a moment in the plane of these two forces (the pitching moment). These forces are functions of the angle of attack of the section.
The lift force a wing section is capable of producing is usually plotted relative to the angle of attack (AOA). See FIGS. 9 and 10 of the drawings. As the wing section AOA increases beyond the stall angle the lift force perpendicular to the AOA decreases abruptly, but the drag force parallel to the AOA increases dramatically and becomes greater than the pre-stall lift force. FIGS. 11 and 12 illustrate this. In FIGS. 11 and 12 L is the lift vector, D is the drag vector, and W is the relative wind direction or AOA. The present invention exploits this post-stall force as will be shown later.
The aerodynamic center (AC) of a wing is defined as the point along the chord line at which the pitching moment is constant, and is at or near the mean quarter chord point, except for delta wings where the AC is at the center of area. Wing lift can be thought of as working through the AC with any pitch moment acting about that pint. Symmetrical wings are wings with no camber (top and bottom surfaces mirror images of each other) and have a zero pitch moment. A wing with positive camber (mean chord line bowed upwardly) will have a negative or downward pitch moment about the AC. The greater the camber the greater the lift force the wing can produce and the greater the negative pitch moment. Wing camber can be varied by using flaps, slats, or ailerons.
Aspect ratio (AR) is defined as the wing span divided by the mean chord length. The aerodynamic characteristics of a wing with a very low AR are much different than for a high AR wing. The lift slope angle is much smaller and it reaches its maximum lift at a much higher angle of attack (AOA) and the maximum lift before stalling is lower.
A delta shaped wing is a low AR wing with additional differences. As the AOA increases a pair of vortex spirals generate at the leading edge and increase in size as they progress to the trailing edge and as the AOA increases. The vortices feed dynamic air over the top surfaces of the wing, lowering the pressure and thus producing lift to angles of attack as high as 80 degrees. Since the AC of a delta wing is already at the center of area there is little movement of the center of lift as the AOA changes. These differences are beneficially exploited by this invention as will be shown later.
The present invention combines a low aspect ratio wing with a delta forebody, which for the sake of brevity will be called the "lifting body", and to which a tail is fixedly attached, and a high aspect ratio wing that is pivotally attached to the lifting body.
Airfoils create lift in two ways. The first way is by making the air travel a longer distance around the foil on the top. This lowers the average air pressure on the top relative to the bottom, and produces lift. This kind of lift is dominant at zero AOA. The other way of creating lift is by deflecting the air downwards (mass transport), much like a rocket throws mass downward creating an equal and opposite reaction. If an airfoil has a positive camber it will create lift at zero AOA and have a negative pitching moment. The pressure distribution of an airfoil sums up to create the net lift and moment effect. Typically the lower side of the foil has a (negative) pressure lower than the free stream pressure because the flow must speed up around the downwardly curving surface. If the lower surface is changed to be made flat for a small distance of say twenty percent (20%) chord, then that surface would not produce negative suction lift buy only positive mass transport lift. Since the bottom surface must converge with the top surface to form an air foil, such a flat angled surface cannot continue very long. By continuing on with a rounded or flat (segmented) surface, the bottom side will converge with the top to form an airfoil. By changing the length of the flat section and its slant into the flow where the downward suction is the greatest, the new foil, with a very similar camber profile as before, can be made to have a POSITIVE pitching moment. This design leads to a higher drag than a normal streamlined airfoil of laminar shape. The bottom surface will be laminar up until the point the flat surface ends, and then be turbulent thereafter. The preferred embodiment of this invention uses flat surfaces on the underside of the lifting body.
Typical laminar foils can achieve laminar flow to seventy percent (70%) of chord on the bottom side. This compares to the loss of around fifty percent (50%) of chord (70-20) that will be turbulent instead of laminar because of the flat surfaces. A laminar foil has typically about fifty-five percent(55%) of the skin in laminar flow, and the rest is turbulent. The flat slanted new airfoil will only have thirty percent (30%) in laminar flow. A laminar flow wing has around 1.45 times more drag than an ideal one hundred percent (100%) laminar flow wing. The new foil would have around 1.7 times that of an one hundred percent (100%) laminar wing. A completely turbulent wing has around 2 times the drag of a one hundred percent (100%) laminar flow wing. Adding in the angle of attack of the slant may, of course, make the new foil have more drag than a turbulent foil, but not my much, if any.
The dominant reasons for having slanted flat surfaces on the bottom of the lifting body is for making water operation practical, and to have a positive pitch moment. A curved surface creates lift, and water is around 800 times as dense as air and therefore creates 800 times as much (negative) lift. A flat surface will not create the negative suction lift but only mass transport positive lift. By having a series of flat surfaces on the bottom surface of the foil that approximate the bottom surface of a normal airfoil, a relatively streamlined shape is achieved. By placing a small step at the point where the flat surface transitions to another flat surface, most of the suction (negative) lift is eliminated up front of the lifting body where the moment arm from the CG is large. With a traditional larger second step near the CG, water operation is practical, and the tendency common to traditional boat hulls and floats, with their curved foreplane surfaces, to plow deeply while getting on the step and to nose over in rough water, is eliminated by the small front step. The aerodynamic and hydrodynamic benefits of the flat surfaces and the two steps more than offset the small increase in drag.
Low aspect ratio or delta type aircraft are limited primarily to high Mach numbers and or high thrust to weight where the low L/D at low speed is not a problem. Conventional aircraft with high aspect ratio wings are limited to lower Mach numbers and smaller angles of attack.
The aircraft of this invention combines the lifting body with a wing that can be pivoted so that the lifting body can reach its optimum angle of attack, near twenty seven (27.degree.) for maximum lift, while the wing remains unstalled at about fifteen degrees (15.degree.) AOA. By having the flaps lowered and the ailerons drooped, all the lifting surfaces are contributing maximum lift for slow speed landings or fast descents.
When the flaps are lowered and the ailerons drooped, the wing camber is increased which increases the negative pitching moment of the wing which must be compensated for to maintain stability. Conventional aircraft increase the negative lift of the tail which increases the load the wing must support. The wing of this invention is located relative to the lifting body such that the AC of the wing is longitudinally behind the AC of the lifting body, as shown in FIG. 2. When the wing is pitched down the incidence angle of the wing becomes negative relative to the incidence angle of the lifting body which produces a positive pitching moment couple that cancels the increased negative pitch moment produced by the concurrent flap deployment. The pitch moments are thus balanced through all wing pitch and flap deployment angles. The aircraft can be flown through the full flight envelope with very little pitch trim and no increased trim drag.
One of the principal advantages of the aircraft invention is the inherent longitudinal stability that results from the AC of the high AR wing being behind the AC of the low AR wing. This can be understood by looking at the lift slope graphs of FIGS. 9 and 10. When the aircraft angle of attack changes due to speed changes, elevator movement, gusts, turbulence, or wind sheer, the lift changes more rapidly on the high AR wing than on the low AR wing. Since the AC of the high AR wing is behind the AC of the low AR wing, the differential lift distribution creates a pitch moment that pitches the aircraft into the relative airstream. This makes the aircraft act like it has a higher static margin (the fore-and-aft distance between the CG and the AC) than it actually has and thus the CG range can be expanded and moved back closer to the AC of the entire aircraft without reducing static and dynamic longitudinal stability. This stability is independent of varying wing pitch angles and flap deployment and is effective through the entire flight envelope.
As a conventional wing stalls the center of lift moves back to the center of area. This coupled with a loss of lift pitches the nose down. If there is enough down force by the tail the nose can be kept up and the plane will mush down with the wing still producing lift. This is very desirable for obvious safety reasons. The aircraft of the invention has this capability without the penalty of high negative tail force. As the wing stalls the aircraft center of lift does not move back to the aircraft center of area. When the wing stalls the lifting body produces relatively more lift than the wing due to the continuing vortex lift of the lifting body. Since the center of lift (COL) of the lifting body is forward of the wing COL the resultant COL shift of the aircraft is small compared to a conventional aircraft. This phenomena plus the CG location near the aircraft AC means the pitch down moment at wing stall is relatively small.
The vortex lift over the lifting body continues after the wing stalls to AOA over about seventy degrees (70.degree.) and the flow remains relatively parallel to the lifting body chord line. This flow continues to feed dynamic air to the tail surfaces at extreme AOA so the aircraft remains controllable in a stalled flight mode. The transition from unstalled to stalled flight and back to unstalled flight is very smooth and requires very little pitch control to maintain attitude. Because of this exceptional control the aircraft is capable of some unusual maneuvers and in an emergency landing can be flared to 60 degrees AOA for extremely low speed at touch down.